The present invention relates generally to gas turbine engines, and, more specifically, to fans and compressors thereof.
In a turbofan gas turbine engine, air is pressurized or compressed in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine which powers the compressor, and also in a following low pressure turbine which powers a fan disposed upstream from the compressor.
A fan is a special form of a compressor having larger rotor blades which pressurize air for providing propulsion thrust for powering an aircraft in flight. The fan defines the first of many compressor stages in which air is increasingly compressed in turn.
Air pressurization is provided by converting rotary energy from the blades into velocity of the air which is then diffused to recover pressure therefrom. Diffusion occurs in locally diverging flowpaths and is limited by undesirable flow separation and corresponding compressor stall.
The fan blades are specifically configured to pump sufficient airflow for generating engine thrust during operation. The multistage compressor is specifically configured to supply high pressure air to the combustor for burning with fuel to generate energy for extraction by the downstream turbines.
A considerable challenge in designing these components is maximizing flow pumping capability and compression efficiency with suitable stall margin, and especially at high thrust conditions where the engine operates closest to its rotational speed and temperature limits. At high rotational speed, the flow Mach numbers relative to the rotor blades are high, and may be supersonic, and the aerodynamic loading or diffusion is also high. The aerodynamic challenge is further complicated by the mechanical and aero-mechanical limitations of the rotor blades themselves.
The fan and the compressor include rotor blades and stator vanes whose airfoils are specifically configured for maximizing performance within conventional constraints. Airfoil design involves many compromises in aerodynamic, mechanical, and aero-mechanical performance. The airfoils have three dimensional (3D) configurations which typically twist in span from root to tip and vary in thickness axially between leading and trailing edges for controlling aerodynamic loading over the corresponding pressure and suction sides thereof.
The flowpath through each compressor stage is defined circumferentially between adjacent blades or vanes, and radially by corresponding outer and inner end walls.
For example, the relatively long fan blades are disposed inside an annular fan casing which defines the radially outer flowpath boundary or outer wall. The blades extend radially outwardly from a supporting disk, and typically discrete inter-blade platforms are suitably mounted to the disk for defining the radially inner flowpath or inner wall.
Similarly, the compressor rotor stages include corresponding rows of rotor blades decreasing in span height in a downstream direction and disposed inside a corresponding annular casing defining a radially outer shroud around each stage. Compressor blades typically include integral blade platforms at the roots thereof which adjoin adjacent platforms for defining the inner flowpath.
And, the corresponding compressor stator stages include vanes affixed at their radially outer ends to an annular outer band typically formed in circumferential or arcuate segments. The radially inner ends of the stator vanes may be plain, or may be affixed to an annular inner band which defines the inner flowpath, which is also typically formed in arcuate segments.
All of the various forms of inner and outer flowpath boundaries described above are similar to each other and are axisymmetric. The outer walls are circumferentially concave and provide a smooth cylindrical or conical surface facing radially inwardly. The inner walls are circumferentially convex and provide a smooth cylindrical or conical surface facing radially outwardly.
For a given engine size and thrust requirement, the sizes of the rotor blades and stator vanes are specified or limited, and cooperate with correspondingly sized outer and inner flowpaths walls. With these deign constraints, the 3D configuration of the blades and vanes is varied in an attempt to maximize flow pumping and compression efficiency with suitable stall margin. Modem three-dimensional viscous computer analysis is used to advantage in designing compressor airfoils, yet performance is nevertheless limited as described above.
Accordingly, it is desired to further improve performance of gas turbine engine compressors and fans within geometric constraints therefor.
A compressor flowpath includes circumferentially spaced apart airfoils having axially spaced apart leading and trailing edges and radially spaced apart outer and inner ends. An outer wall bridges the airfoil outer ends, and an inner wall bridges the inner ends. One of the walls includes a flute adjacent the leading edges for locally increasing flow area thereat.